Method of manufacturing a structure

ABSTRACT

A method of manufacturing a structure, the method comprising: forming a panel assembly by bonding a stack of thickness control plies of composite material to a laminar composite panel, the stack of thickness control plies having a first edge proximate an edge of the laminar composite panel, a second edge opposite the first edge, and a ramp where the thickness of the stack of thickness control plies decreases towards the first or second edge; measuring the thickness of the laminate composite panel or the panel assembly; controlling the number of thickness control plies in accordance with the measured thickness; attaching a first component to the panel assembly, the first component having a surface which engages the laminar composite panel and a ramp which engages the ramp in the stack of thickness control plies; and attaching the laminar composite panel at or near its edge to a further component.

RELATED APPLICATION

The present application is based on, and claims priority from, BritishApplication Number 0910938.0, filed Jun. 25, 2009, the disclosure ofwhich is hereby incorporated by reference herein in its entirety.

FIELD OF THE INVENTION

The present invention relates to a method of manufacturing a structurein which a laminar composite panel is attached to a first component, andthe laminar composite panel is attached to a further component after ithas been attached to the first component. By way of non-limiting examplethe panel may be a cover of an aircraft wing, the first component may bea spar of the wing, and the further component may be a centre wing box.

BACKGROUND OF THE INVENTION

FIG. 1 is a schematic cross sectional view of an aircraft wing rootjoint comprising an aircraft wing box 1 attached to a centre wing box10. The centre wing box 10 comprises a rib 11 connecting upper and lowerroot joint fittings 2, 3. The aircraft wing box 1 comprises upper andlower laminar composite covers 4, 5 connected by a spar 6 which extendsbetween them.

The covers 4, 5 have inner and outer mould lines 4 a, 5 a, 4 b, 5 b. Theouter mould lines 4 b, 5 b form external aerodynamic surfaces of thewing box 1, while the inner mould lines 4 a, 5 a, form internal surfacesof the wing box 1 opposite the outer mould lines 4 b, 5 b.

The spar 6 is shown in longitudinal section in FIG. 1. The spar 6 isC-shaped in transverse section with upper and lower spar flanges 7, 8which extend from a web 9 and engage the inner mould lines 4 a, 5 a ofthe covers 4, 5. The spar flanges 7, 8 terminate outboard of the inboardedge of the spar web 9 to leave spar flange cutback areas where thecovers 4, 5 overhang the flanges 7, 8 in an inboard span-wise directiontowards the root joint fittings 2, 3.

A problem with manufacturing such a joint is that the thickness of thelaminar composite covers 4, 5 cannot be controlled accurately due to themethod of manufacturing the laminate and the properties of the compositematerial. As a result the covers 4, 5 may not precisely meet the rootjoint fittings 2, 3 resulting in either a gap or a clash. This problemcan be solved by bending the covers 4, 5 (where they overhang the sparflanges 7, 8) either towards the fittings (to close a gap) or away fromthe fittings (to avoid a clash). However this will induce stress in thecovers 4, 5 and as a result the covers must be made thicker and heavierto cope with this stress. Also it may be difficult to bend the coverssufficiently, particularly near the cover/spar interface shown in FIG.1.

The thickness of the upper cover 4 (i.e. the distance between the innermould line 4 a and the outer mould line 4 b) is particularly critical asit is its outer mould line 4 b (rather than its inner mould line 4 a)which engages the root joint fitting 2. Clearly, any increase ordecrease in the thickness of the upper cover 4 will affect the alignmentbetween the upper cover 4 and the upper root joint fitting 2 and/or thealignment between the lower cover 5 and the lower root joint fitting 3.Conversely, as it is the inner mould line 5 a of the lower cover 5 whichengages the lower root joint fitting 3 (rather than its outer mould line5 b), the thickness of the lower cover does not affect the alignmentbetween the lower root joint fitting 3 and the lower cover 5, nor doesthe thickness of the lower cover affect the alignment between the uppercover 4 and the upper root joint fitting 2. However, the thickness ofthe lower cover is critical for the attachment of a buttstrap 12 whichconnects the lower cover OML 5 b to a lower cover 13 of the centre wingbox 10. Therefore, any thickness variation in the manufacturing processof both covers must be carefully managed.

SUMMARY OF THE INVENTION

A first aspect of the invention provides a method of manufacturing astructure, the method comprising

-   -   a) forming a panel assembly by bonding a stack of thickness        control plies of composite material to a laminar composite        panel, the stack of thickness control plies having a first edge        proximate an edge of the laminar composite panel, a second edge        opposite the first edge, and a ramp where the thickness of the        stack of thickness control plies decreases towards the first or        second edge;    -   b) measuring the thickness of the laminate composite panel or        the panel assembly;    -   c) controlling the number of thickness control plies in        accordance with the measured thickness;    -   d) attaching a first component to the panel assembly, the first        component having a surface which engages the laminar composite        panel and a ramp which engages the ramp in the stack of        thickness control plies; and    -   e) attaching the laminar composite panel at or near its edge to        a further component.

The various steps of the method may be performed in any order, andindeed some of the steps (for instance steps a), d) and e)) may beperformed at the same time, but preferably step e) is performed afterstep d).

A second aspect of the invention provides a structure comprising:

-   -   a panel assembly comprising a stack of thickness control plies        of composite material bonded to a laminar composite panel, the        stack of thickness control plies having a first edge proximate        an edge of the laminar composite panel, a second edge opposite        the first edge, and a ramp where the thickness of the stack of        thickness control plies decreases towards the first or second        edge;    -   a first component attached to the panel assembly, the first        component having a surface which engages the laminar composite        panel and a ramp which engages the ramp in the stack of        thickness control plies; and    -   a further component attached to the laminar composite panel at        or near its edge.

The laminar composite panel may be attached at or near its edge to thefurther component by an overlapping joint, a butt joint, or any othersuitable connection.

Typically a longest one of the thickness control plies extends alongonly part of a length of the laminar composite panel, preferably lessthan 50% of its length.

Preferably the thickness of the stack of thickness control pliesdecreases towards the second edge.

In the event that the laminate composite panel is too thick or thin,then it may be bent to form a ramp after step c). The ramp may beinclined in the opposite sense to the ramp in the stack of thicknesscontrol plies, or in the same sense.

Typically at least some of the plies of the laminar composite panel aremade from Carbon Fibre Reinforced Plastic (CFRP), although othercomposite materials may be envisaged such as Glass-Fibre ReinforcedPlastic.

Further preferred features of the present invention are set out in thedependent claims.

In the embodiment of the invention described the further component is acentre wing box of an aircraft wing, the first component is a spar of anaircraft wing and the panel assembly is an aircraft wing cover assembly.However the invention is applicable to a variety of other joints. Forinstance the further component may be an aircraft fuselage skin panel,the first component may be a frame of an aircraft fuselage, and thepanel assembly may be an aircraft fuselage skin panel assembly which isattached to the first component by a butt-strap.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the invention will now be described with reference to theaccompanying drawings, in which:

FIG. 1 is a cross-sectional schematic view of an aircraft root wingjoint;

FIGS. 2 and 3 illustrate a method of managing the thickness of acomposite panel;

FIG. 4 is a cross-sectional schematic view of an aircraft root wingjoint where the thicknesses of the wing box covers have been managedusing the method illustrated by FIGS. 2, 3;

FIG. 5 is a perspective view of the wing box shown in FIG. 4;

FIG. 6 is a cross-sectional schematic view of a cover to spar interfacewhere the cover has a thickness at the lower end of a manufacturingtolerance; and

FIG. 7 is a similar view to FIG. 6 but where the cover has a thicknessat the upper end of a manufacturing tolerance.

DETAILED DESCRIPTION OF EMBODIMENT(S)

FIGS. 2 a and 2 b show a pair of partially cured Carbon Fibre ReinforcedPlastic (CFRP) aircraft wing covers 20, 22 on a mould tool 24. Thecovers 20, 22 have outer mould lines (OMLs) 21, 23, which engage themould tool 24, and non-moulded inner mould lines (IMLs) 25, 27 oppositethe outer mould lines 21, 23. The covers are moulded by laying a vacuumbag (not shown) over the IMLs, applying a vacuum between the vacuum bagand the covers to press the covers against the mould tool 24, andheating the covers to cure them.

The thicknesses of the covers 20, 22 are not accurately controllable andcan range from a lower tolerance value X (see FIG. 2 a) to an uppertolerance value Y (see FIG. 2 b). As explained above with respect toFIG. 1, these relatively large uncertainties in cover thickness areproblematic at the aircraft wing root joint.

In order to manage the cover thickness tolerances in the region of theroot joint, a stack 26 of uncured thickness control plies is laid up onthe IMLs of the covers 20, 22 and the cover/thickness control plyassemblies are either co-cured, or the thickness control plies areco-bonded onto the cover during a further curing cycle. In all cases,the stack 26 has a ramped section 32 with a ramp 33 (that is, aninclined surface) where the thickness of the stack 26 decreases towardsits outboard edge, and a flat section 34 towards its inboard edge. Theramped section 32 may be formed, for example, by laying up progressivelyshorter thickness control plies on top of one another or by selectivelydiscontinuing internal plies. Referring to FIGS. 3 a and 3 b, thethickness of the cured panel assembly is first measured, and then ifnecessary some of the thickness control plies are removed from the stack26 by a machining process. The number of plies which are removed willthus depend on the measured thickness. This results in machined andcured IML surfaces 36, 38 which are a known distance D from the OMLs 21,23. The distance D is subject only to the tolerances of the machiningand measuring processes, regardless of the thickness of the covers 20,22. The tolerances involved in machining and measuring are typically farless than those involved in manufacturing CFRP covers 20, 22. Therefore,the distance D can be controlled with far greater accuracy than thethickness of the laminar composite covers 20, 22. Note that the stack 26of thickness control plies extends along only a relatively small portionof the span of the wing close to its inboard edge, the span-wise lengthof the thickness control plies being determined by the stress allowableof the covers. This minimises both the size of the required machiningtool and the weight of the added thickness control plies.

FIG. 4 shows a cross sectional schematic view of the root of an aircraftwing box 39. The wing box comprises upper and lower laminar compositecovers 40, 41 which are bolted to a spar 42 which extends between them.Note that only the inboard root end of the spar and wing box is shown inFIG. 4. The spar 42 comprises upper and lower spar flanges 44, 45extending from a spar web 46, and the flanges 44, 45 engage the upperand lower covers respectively. The spar flanges terminate outboard ofthe inboard edge of the spar web 46 to leave spar flange cutback areaswhere the covers overhang the flanges in an inboard span-wise direction.The covers 40, 41 have been manufactured by the process described abovewith reference to FIGS. 2 a-3 b. As the upper and lower covers (and theupper and lower spar flanges) are similar, only the upper cover (and theupper spar flange) will be described in detail.

The upper cover 40 has a moulded surface (OML) 48, a non-moulded surface(IML) 50 opposite the OML 48 and a stack 52 of thickness control pliesco-bonded or co-cured to the IML 50 of the cover. The thickness of thestack 52 control plies has been controlled as described above to form acontrolled EVIL surface 53 which is a precise distance D from the OML48. The stack 52 consists of a ramped outboard section 54 with a ramp 55where the thickness of the stack 52 decreases towards its relativelythin outboard edge, and a flat inboard section 56 which terminates atthe same span-wise position as the spar flange 44. The spar web 46converges at its inboard end, and the flange 44 is joggled to follow theline of the spar web and form a ramped surface 58 and a flat un-rampedsurface 59. The ramp 58 in the spar flange engages the ramp 55 in thestack, and the un-ramped surface 59 of the spar flange engages themachined surface 53 of the stack 52. The covers and spar flanges arebolted together.

Still referring to FIG. 4, the upper and lower covers 40, 41 are alsobolted to respective cruciform and tri-form root joint fittings 65, 67of a centre wing box. The spar web 46 is attached to a spar web (notshown) of the centre wing box via a vertical buttstrap (also not shown).

The centre wing box also has a rib 68 (conventionally referred to as“rib one”) connecting the upper and lower root joint fittings 65, 67. Asthe distance D between the inner and outer mould lines of the uppercover 40 can be controlled with close accuracy, the problem ofgaps/clashes arising between the wing covers 40, 41 and the root jointfittings 65, 67 is greatly reduced (or even eliminated). As a result,the covers do not need to be bent significantly into shape duringinstallation of the root joint, which means that fewer stresses andstrains need to be tolerated by the covers. Thus, the covers 40, 41 canbe made thinner and lighter. Similarly, as the distance D between theinner and outer mould lines of the lower cover can be closelycontrolled, the buttstrap 69 can be attached with less measuring andmachining, thus improving the quality and reducing assembly time.

In the case illustrated in FIG. 4 (referring to the upper cover 40 butalso applicable to the lower cover 41) the thickness of the cover isequal to its ‘design value’—that is, the cover has a thickness whichlies in the centre of a manufacturing tolerance range. Consequently, theouter mould line 48 forms a continuous flat aerodynamic surface, even inthe ramped region.

FIG. 5 is a perspective view of the wing box 39 taken from above withthe upper cover removed. FIG. 5 also shows ribs 61, 62 (conventionallyreferred to as “rib two” and “rib three” respectively) which connect thefore and aft spars of the wing box 39 as well as the upper and lowercovers. The ribs 61, 62 are bolted to the upper and lower covers and tothe webs 46 of the fore and aft spars.

FIG. 6 is similar to FIG. 4 and identical features are given the samereference numerals. Note that the lower cover has been omitted forclarity. In this case, the thickness of the cover 40 a is equal to thelower tolerance value X (see FIG. 2 a). In order to achieve the requiredvalue of D, the thickness of the stack 52 a of thickness control pliesis greater than the thickness of the (machined) stack 52 shown in FIG.4. When the cover 40 a and the stack 52 a of thickness control plies arebolted to the spar, the portion of the cover outboard of the stack 52 ais bent down towards the upper spar flange 44. This creates a joggledsection 70 a in the cover 40 a which is inclined in the opposite senseto the ramp 55 in the stack of thickness control plies.

The angle of inclination of the ramp in the joggled section 70 a istypically very shallow (i.e. within the bending allowables) and haslittle effect on the aerodynamic performance of the wing. Note that theangle of the joggled section 70 a shown in FIG. 6 is exaggerated forclarity.

FIG. 7 is similar to FIGS. 4, 6 and identical features will be given thesame reference numerals. Note that the lower cover has again beenomitted for clarity. In this case the thickness of the cover 40 b isequal to the upper tolerance value Y (illustrated in FIG. 2 b). When thecover 40 b and the stack 52 b of thickness control plies are bolted tothe spar, the portion of the cover outboard of the stack 52 b is bent upto create a joggled section 70 b with a ramp which is inclined in thesame sense as the ramp in the stack of thickness control plies. The rampin the joggled cover engages the top of the ramp in the spar flange.

The ramp in the joggled section 70 b is typically very shallow (i.e.within the bending allowables) and has little effect on the aerodynamicperformance of the wing. Again, no shimming is required, which savestime and complexity. Note that, again, the angle of the joggled section70 b shown in FIG. 7 is exaggerated for clarity. Note also that thejoggle has a downward inboard incline as opposed to the upward inboardincline of the joggled section 70 a shown in FIG. 6.

As noted above, in each case the stacks 52, 52 a, 52 b decrease inthickness towards their outboard edges. Alternatively, the stacks 52, 52a, 52 b may decrease in thickness towards their inboard edges (i.e. theedges proximate the inboard edges of the covers 40, 40 a, 40 b). In thiscase, the height of the spar web would be matched accordingly with thestacks.

In the method described above, the thickness of the cured panel assemblyis first measured, and then if necessary some of the thickness controlplies are removed from the stack 26 by a machining process. In analternative case in which the thickness control plies are bonded to apreviously cured cover, then the thickness of the cured cover may bemeasured before the thickness control plies are bonded in place, and thenumber of thickness control plies selected in accordance with themeasured thickness (either by peeling off some of the uncured plies in apre-prepared stack, or by laying up the stack with exactly the rightnumber of plies). Thus in this case it is not necessary to machine awaysome of the cured thickness control plies after they have been bonded tothe cover.

Although the invention has been described above with reference to one ormore preferred embodiments, it will be appreciated that various changesor modifications may be made without departing from the scope of theinvention as defined in the appended claims.

The invention claimed is:
 1. A method of manufacturing a structure, themethod comprising: a) forming a panel assembly by bonding a stack ofthickness control plies of composite material to a laminar compositepanel, the stack of thickness control plies having a first edgeproximate an edge of the laminar composite panel, a second edge oppositethe first edge, and a ramp where the thickness of the stack of thicknesscontrol plies decreases towards the first or second edge; b) measuringthe thickness of the laminate composite panel or the panel assembly; c)controlling the number of thickness control plies in accordance with themeasured thickness; d) attaching a first component to the panelassembly, the first component having a surface which engages the laminarcomposite panel and a ramp which engages the ramp in the stack ofthickness control plies; and e) attaching the laminar composite panel ator near its edge to a further component.
 2. The method of claim 1wherein the number of thickness control plies is controlled in step c)by removing one or more of the plies from the stack of thickness controlplies either before or after the stack of thickness control plies hasbeen bonded to the laminar composite panel in step a).
 3. The method ofclaim 2 wherein the thickness of the panel assembly is measured in stepb), and the number of thickness control plies is controlled in step c)by removing one or more of the plies from the stack of thickness controlplies after the stack of thickness control plies has been bonded to thelaminar composite panel in step a).
 4. The method of claim 1 wherein thethickness of the laminar composite panel is measured in step b), and thenumber of thickness control plies is controlled in step c) by selectingthe number of plies in the stack of thickness control plies before it isbonded to the laminar composite panel in step a).
 5. The method of claim1 wherein the thickness of the stack of thickness control pliesdecreases towards the second edge.
 6. The method of claim 1 furthercomprising bending the laminar composite panel after step a) to form aramp in the laminar composite panel.
 7. The method of claim 6 whereinthe ramp is inclined in the opposite sense to the ramp in the stack ofthickness control plies.
 8. The method of claim 6 wherein the ramp inthe laminar composite panel is inclined in the same sense as the ramp inthe stack of thickness control plies, and wherein the ramp in thelaminar composite panel engages the ramp in the first component.